Liquid propellant rocket engine coaxial injector

ABSTRACT

To provide combustion stability in rocket engines over a wide range of thrust, a coaxial injector is provided which includes an outer annular sleeve for introduction of one propellant and a hollow pintle with spaced slots for introduction of another propellant to impinge on the first propellant. The slots are staggered in location and size to permit initial interlocking of the propellants and, therefore, increased mixing and combustion performance.

[ 51 Oct. 24, 1972 Seamans, Jr..........60/39.74 A 3,462,950 8/1969Chevalaz ..............60/39 74 A [54] LIQUID PROPELLANT ROCKET ENGINECOAXIAL INJECTOR [72] Inventor: Gerard W. Elverum, Jr., Rolling Pn-maryExaminer samuel Feinberg Hl Callf- Attorney-George C. Thompson Toprovide combustion stability in rocket engines over a wide range ofthrust, a coaxial injector is provided which includes an outer annularsleeve for introduction of one propellant and a hollow pintle withspaced slots for introduction of another propellant to impinge on thefirst propellant. The slots are staggered in location and size to permitinitial interlocking of the propellants and, therefore, increased mixingand combustion performance.

T M m MI 5 50 4 20 6 K MU r C m h 8 1% C 5 a 2 ./5 e "O5 B O "6 6 u .8.m nm4 m A m m 4 a 6 w m mm R 8 m" m m v6 lnus 9 0 u "2 m .2 m 6 mmh 9 2m h 6 mmmM u O "n W N "f n CID .W 1 d 8 e D. 8.!- S "H 0- .mm A F A UIFMMMMH NUT 7 2 2 555 .l.[[ [ill 1 Claim, 3 Drawing Figures [56]References Cited UNITED STATES PATENTS 3,093,157 6/1963 Aitken et al.................60/258 PATENTEDnm 24 I972 SHEET 1 BF 2 Gerard W.Elverum, Jr.,

INVENTOR.

ATTORNEY.

PATENTED um 24 m2 SHEET 2 BF 2 Gerard W. Elverum, Jr,

Fig. 5.

INVENTOR.

ATTORNEY.

LIQUID PROPELLANT ROCKET ENGINE COAXIAL INJECTOR BACKGROUND OF THEINVENTION This invention pertains to rocket engines and moreparticularly to a simple injector concept for interlocking two or morepropellants to provide increased combustion efficiency and to provide afundamental mechanism for achieving dynamically stable combustion in awide range of engine sizes.

One of the greatest problems in rocket engines when two propellants areintroduced through an injector for mixing, is that of achievingcombustion stability. This problem of maintaining combustion stabilityhas been present or can be initiated in nearly all rocket engines, andparticularly as the size of the engine is increased. By using aninjector concept in which all of the propellant is introduced into thechamber from the center region of the head end closure plate, thecombustion process can be made dynamically stable in all engine sizes.This results from the energy source being located at the nodal point ofthe acoustic pressure field. However, a fundamental problem exists,particularly with hypergolic propellants, when the injection into thechamber is concentrated at the center of the engine. This problem withhypergolic propellant combustion is the degree of hydraulic mixing whichdetermines both the propellant distribution and the heat available forvaporization. This mixing becomes increasingly limited as the injectorelement size is increased. As larger streams or solid sheets of veryreactive oxidizer and fuel impinge on each other, prior art enginesexperienced only limited liquid mixing before gas and vapor generated atthe innerface is present in sufficient quantity to partially separatethe oxidizer and fuel streams. The result of this is that a certainpercentage of the total propellant is deflected apart and remainsunmixed, and any further mixing must occur downstream in the combustionchamber either by the relatively slow processes of diffusion orre-circulation. This results in decreased combustion efficiency in theengine. In addition to the problem of not obtaining sufficient mixingwith larger element sizes, the concentration of the propellant intorelatively unmixed pockets can cause high amplitude local pressurefluctuations if they reach explosion limits. These fluctuations maybecome quite severe in the fuel-rich regions of the chamber, and in somecases large enough pressure spikes to be destructive of the rocketengine walls can occur. In addition, these disturbances form the mainsource of triggers for combustion instability in the chamber.

In prior art structures, the problem of combustion instability has beenresolved in some cases by accepting the possibility of the phenomena andstatistically assuming that a certain percent of the engines will fail.Another way of solving this problem has been to provide baffles on theinjector face or other types of acoustic energy absorbing devices in thecombustion chamber to decrease the tangential or transverse and radialmodes of pressure fluctuations. However, with these methods, for eachchange in engine size, much design and testing work is necessary toprovide the proper baffle or absorber configuration and also toestablish the other injector design parameters to provide a matchedsystem. An example of this baffle construction is seen in the US. Patentto Mower et al., No. 3,200,589.

SUMMARY OF THE INVENTION This invention obviates the problems andcomplexities of prior art structures by providing a single, centrallylocated coaxial injector with one propellant impinging on the other inan annular stream. The other propellant is introduced throughgeometrically staggered orifices having several different areasto-provide penetration and interlocking of the oxidizer and fuel in theimpingement zone prior to the occurrence of significant liquid and/orvapor phase reactions.

BRIEF DESCRIPTION OF THE DRAWINGS FIG. 1 is a perspective view,partially in cross section, of a coaxial injector according to thisinvention;

FIG. 2 is a view in cross section of aninjector according'to thisinvention; and

FIG. 3 illustrates the arrangement of slots in the pintle forintroduction of one of the propellants.

DESCRIPTION OF THE PREFERRED EMBODIMENTS Referring to FIG. 1, aninjector constructed according to this invention is shown partially incross section. The coaxial injector is shown generally at 2 and extendsinto the combustion chamber 4 whichis formed or enclosed by combustionchamber wall 6. An outer annular sleeve 8 forms with hollow pintle 1 0an annular orifice or slot 12 for introduction of propellant such asfuel in the direction of arrows 13. A closure or cap member 14 isattached to or integral with pintle 10. A first series of slots 16alternate with a plurality of smaller slots 18 to allow propellant suchas oxidizer introduced to the'interior of pintle 10 to exit in a radialdirection as shown. by arrows 20, to penetrate, interlock andhydraulically mix with the propellant from annular orifice or slot 12. Acone shaped projection 22 diverts the propellant toward slots l6and 18.In FIG. 2, the fuel ducts 26 are shown as illustrating one manner inwhich propellant is supplied to orifice or slot 12 through orifices 27and manifold 29. Oxidizer enters at 28. A central member 30 is shownwith supporting vanes 32 and 34 to provide support for the pintle tip.In a similar manner, vanes 35 are provided to properly space pintle 10from sleeve 8. FIG. 3 is illustrative of the slots 16 and 18. In thisconnection, while slots 18 are shown as resembling a keyhole, it iswithin this inventions scope to provide rectangular or other shapedslots. Similarly, the shape of slots l6can be varied.

It has been found by applicant that the area of slots 16 relative toslots 18 are selected based on the propellant flow rate to pintlediameter ratio. The optimum range for the percentage of propellant whichshould flow through openings 16 lies between 5 and 50 percent. Inaddition, the unit spacing between the slots 18 is such as to bedirectly proportional to the thickness of the propellant sheet exitingorifice 12 such that radial and circumferential hydraulic interlockingof the two propellants is proportionately the same at all engine sizes.

' In operation, propellant such as fuel is injected through orifice 12in a substantially axial direction.

Propellant such as oxidizer exits in a radial direction through thestaggered slots 16 and 18 to interlock and impinge with the fuel. Byproviding the small slot 18 between the main slots 16, an increase inefficiency is provided by additional mixing and reaction of thepropellants which have been caused to bypass the larger slots. Thishydraulic interlocking prior to reaction permits all of the propellantsto be introduced at the center portion of the combustion chamber withoutattendant loss in combustion efficiency even in very large engine's.

By this described configuration using a coaxial injection with the slotarrangement as shown and described, it has been found that regardless ofengine size the problem of combustion instability no longer exists.

The damaging tangential modes of combustion instability which may beinitiated by a disturbance in the combustion chamber or feed system isimmediately dynamically damped by this central injection method. Inaddition, the first radial mode of instability is also effectivelydamped by this configuration.

Having described this invention, it is understood that it is to belimited only by the scope of the claims appended hereto.

I claim:

1. In a gas generating device having a combustion chamber, an injectorand a throat area for exit of gas to produce thrust, that improvement insaid injector which comprises:

an outer annular sleeve member,

an inner substantially hollow member, said outer and inner membersforming an annular orifice for the introduction of a first propellantinto said combustion chamber,

said inner member having one end extending into said combustion chamberand having a closure substantially covering one end of said innermember, said one end having a plurality of first orifices adapted toallow a second propellant to interlock and mix with said firstpropellant,

said one end further having a plurality of second orifices, each of saidsecond orifices being circumferentially and axially spaced from saidfirst orifices and said second orifices each having a crosssectionalarea smaller than each of said first orifices wherein circumferentialspacing of said first orifices is proportional to the thickness of saidannular orifice.

1. In a gas generating device having a combustion chamber, an injectorand a throat area for exit of gas to produce thrust, that improvement insaid injector which comprises: an outer annular sleeve member, an innersubstantially hollow member, said outer and inner members forming anannular orifice for the introduction of a first propellant into saidcombustion chamber, said inner member having one end extending into saidcombustion chamber and having a closure substantially covering one endof said inner member, said one end having a plurality of first orificesadapted to allow a second propellant to interlock and mix with saidfirst propellant, said one end further having a plurality of secondorifices, each of said second orifices being circumferentially andaxially spaced from said first orifices and said second orifices eachhaving a cross-sectional area smaller than each of said first orificeswherein circumferential spacing of said first orifices is proportionalto the thickness of said annular orifice.